Gas turbine engines are known to have marginal combustion at very high altitudes which is attributable at least in part to the low fuel flows per injector In fact, fuel flows at high altitudes on the order of fifty thousand feet are frequently quite low during starting, e.g., less than three pounds per hour per injector. Moreover, the high fuel viscosity encountered in cold high altitude conditions adds further difficulty to achieving reliable starting.
In instances where, after flameout, a restart must be achieved at high altitude, then special means are often employed. Hence pyrophoric fuels, which ignite spontaneously on contact with air, can be employed during starting. These are dangerous chemicals, typically restricted to use in military aircraft, which are regarded as undesirable.
If the gas turbine engine experiences flameout, at high altitude it is usually necessary to descend to a much lower altitude before it can be restarted Typically, this may require a descent to an altitude of thirty thousand feet or less which is most undesirable in combat and since the gas turbine engine will most likely spool down. Once engine spool down occurs, it is most difficult to achieve high altitude starting where conventional liquid fuel is utilized.
Additionally, the starting of a gas turbine is made difficult because of problems in providing adequate fuel atomization even at altitudes on the order of thirty thousand feet. At higher altitudes, starting is also limited by kinetic loading. Additionally, even at full engine speed kinetic loading limits, i.e., difficulty in burning or completing the combustion reaction, is limiting and it causes a tendency for flameout and combustion inefficiency.
The present invention is directed to overcoming one or more of the above problems.